Supersonic aircraft with spike for controlling and reducing sonic boom

ABSTRACT

Method and arrangement for reducing the effects of a sonic boom created by an aerospace vehicle when the vehicle is flown at supersonic speed. The method includes providing the aerospace vehicle with a first spike extending from the nose thereof substantially in the direction of normal flight of the aerospace vehicle, the first spike having a second section aft of a first section that is aft of a leading end portion, the first and second sections having a second transition region there between and each of the sections having different cross-sectional areas, the leading end portion of the first spike tapering toward a predetermined cross-section with a first transition region between the predetermined cross-section and the first section. The first transition region is configured so as to reduce the coalescence of shock waves produced by the first spike during normal supersonic flight of the aerospace vehicle.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application is a divisional application of U.S. applicationSer. No. 11/754,276, filed 26 May 2007, now issued as U.S. Pat. No.8,083,171, which is a continuation application of U.S. application Ser.No. 11/307,280, filed 30 Jan. 2006, now abandoned, which is a divisionalapplication of U.S. application Ser. No. 10/708,404, filed 1 Mar. 2004,now abandoned, which is a continuation-in-part application ofPCT/US03/02631, filed 30 Jan. 2003, and a continuation-in-partapplication of U.S. application Ser. No. 10/104,403, filed 22 Mar. 2002and issued as U.S. Pat. No. 6,698,684, each of which are acontinuation-in-part application of U.S. application Ser. No.10/060,656, filed 30 Jan. 2002, now abandoned. Said applications arehereby expressly incorporated by reference into the present applicationin their entireties for purposes of disclosure.

TECHNICAL FIELD

The invention relates generally to supersonic aircraft fuselage design.More particularly, it relates to aircraft accessory designs forcontrolling the magnitude of pressure disturbances or waves generated byan aircraft flying at supersonic speed so as to reduce sonic boomeffects at ground level.

BACKGROUND ART

In flight, an aircraft produces pressure waves or disturbances in theair through which it is flying. These pressure waves propagate at thespeed of sound. When the aircraft flies at subsonic speed, thesepressure waves propagate in all directions around the aircraft,including ahead of the aircraft. When aircraft fly at supersonic speed,these pressure waves cannot propagate ahead of the aircraft because theaircraft is traveling faster than the propagation speed of the waves.Instead, the pressure waves generated by the aircraft typically coalesceinto two shock waves, one associated with the nose of the aircraft andthe other associated with the tail of the aircraft. These shock wavespressure differentials that propagate circumferentially away from theaircraft. With respect to the shock wave associated with the nose (the“bow shock”), the pressure increases abruptly from about ambient to apressure significantly thereabove. The pressure decreases down from thispressure significantly above ambient down to a pressure below ambient inthe region between the bow shock and the shock wave associated with thetail (the “tail shock”). The pressure then increases abruptly from belowambient to about ambient at the tail shock.

These shock waves can propagate great distances away from the aircraftand eventually those that are directed downwardly will reach the groundwhere they can produce significant acoustic disturbances called sonicbooms. Sonic booms are so named because of the sounds created by theabrupt pressure changes when the shock waves pass a reference point onthe ground. The acoustic signature of a sonic boom is commonlycharacterized as an N-wave because the pressure changes associated withthe acoustic signature resemble the letter “N” when plotted as afunction of position from the nose of the aircraft. That is, an N-waveis characterized by the abrupt pressure rise associated with the bowshock, commonly referred to as “peak overpressure,” followed by adecrease to a pressure below ambient. This is followed by the abruptrise back toward ambient pressure in association with the tail shock.Where perceivable, typically on the ground by a person, a sonic boomeffect is caused by the two rapid-succession, high magnitude pressurechanges. Strong sonic booms cause an objectionably loud noise, as wellas other undesirable conditions at ground level. For these reasons,supersonic flight over some populated areas is restricted. A schematicrepresentation of the phenomenon of aircraft produced sonic boom isprovided in FIG. 20.

It should be appreciated that shock waves propagate in the form of a“Mach Cone” having a shape defined by a Mach angle (μ). The Mach angle μis a function of the Mach number M, which is defined as the ratio of thespeed of an object over the speed of sound. The Mach angle (μ.) can bedetermined using the equation:sin(μ)=1/M, orμ=sin.⁻¹(1/M)

The shape of the Mach cone produced by an aircraft in supersonic flightcan be represented by rotating a line drawn from the aircraft's nose tiptoward the tail of the aircraft and oriented at an angle (μ) withrespect to the aircraft's direction of travel. Consequently, the tip ofthe Mach cone points in the direction of travel.

In order for supersonic flight over land to be acceptable, the pressuredisturbances that cause the sonic boom's acoustic signature must becontrolled to avoid strong sonic boom effects caused by the abruptpressure changes due to the bow and tail shock waves.

It should be appreciated that it is not only the magnitude of thecreated pressures that are radiated to ground level from an aircraftflying at supersonic speeds that causes persons to experience unpleasantsonic boom effects, but it is primarily the rate(s) of change in thepressures experienced at ground level (pressure differentials—ΔP) thatproduces the undesirable sonic boom effects. Therefore, one goal forminimizing audible sonic boom effects is to control pressuredifferentials caused at ground level by a supersonic flying craft.

Another characteristic of the pressure waves or disturbances generatedby a supersonic flying aircraft is that the elevated pressuresassociated essentially with the forward portion of the craft have aneffect that coalesces together as they travel toward the ground. As FIG.20 depicts, the lowered pressures associated essentially with therearward portion of the craft also have an effect that coalescestogether as they travel toward the ground. As described above, it isthese two primary pressure changes that cause the sonic boom effects atground level. Therefore, it can be a solution to the sonic boom problemto smooth the pressure differentials so that there are no abruptchanges. That is to say, the magnitude of the different pressuresinduced by a supersonic flying aircraft need not necessarily be altered,but it can be enough for some aircraft designs to smooth the abruptpressure changes experienced at ground level to be more gradual.

Features of the aircraft that cause such abrupt changes in the inducedpressures are also detrimental. As explained hereinabove, the pressuredisturbances or waves radiate from the aircraft at a relationship basedat least in part on the speed of the craft. The angle of radiation canalso be affected by the magnitude of the caused disturbance. That is tosay, and is best illustrated in FIG. 21, abrupt projections off of thefuselage of the aircraft (transverse to the direction of travel of theaircraft) will cause larger and higher angle disturbances than smoothtransitions. In the case of FIG. 21, the outwardly projecting jetengines cause pressure waves; one at the top, forward projecting portionof the inlet, and another at the lower lip of the engine's inlet. Thepressure disturbances induced by the engine of the aircraft in FIG. 21coalesce and thereby detrimentally create a combined pressuredifferential at the ground. Therefore, working toward the goal ofminimizing differentials in the pressure profile or signature of asupersonic aircraft, a design challenge has been identified to keeptransverse projections (to the direction of travel of the aircraft), andeven surface disruptions to a minimum. In this context, a surfacedisruption is considered to be any dimensional change along the lengthof the aircraft that is transverse to the axis of travel. Since it ispressure waves radiating from the bottom of the plane that most effectsground boom, it is to the extreme lower surfaces of the aircraft thatthis smoothing goal is most relevant.

As background to the present invention(s), it is known that attemptshave been made to modify the design of supersonic aircraft in order toadjust the sonic boom signature. These modifications have includedchanges to wing design, as described in U.S. Pat. No. 5,934,607, issuedto Rising, et al., for a “Shock Suppression Supersonic Aircraft.”Another approach involves incorporating air passages through thefuselage or wings of supersonic aircraft, such as the structuresdescribed in U.S. Pat. No. 4,114,836, issued to Graham, et al., for an“Airplane Configuration Design for the Simultaneous Reduction of Dragand Sonic Boom”; U.S. Pat. No. 3,794,274, issued to Eknes, for an“Aircraft Structure to Reduce Sonic Boom Intensity”; and U.S. Pat. No.3,776,489, issued to Wen, et al., for a “Sonic Boom Eliminator.” Furtherattempts at reducing the sonic boom caused by supersonic aircraftinclude the addition to the aircraft of structure arranged to disruptthe air flow patterns as the aircraft travels at supersonic speed.Examples include the structure described in U.S. Pat. No. 3,709,446,issued to Espy, for a “Sonic Boom Reduction” and U.S. Pat. No.3,647,160, issued to Alperin, for a “Method and Apparatus for ReducingSonic Booms.”

Another attempt to control the sonic boom in a supersonic aircraft usesa blunt nose to increase the air pressure immediately adjacent to thenose of the aircraft, thus disrupting the normal formation of thepressure wave that causes the acoustic signature. This disruptionresults in a reduction of the abruptness of the pressure changes thatdevelop after the initial pressure rise in the acoustic wave thatstrikes the ground. A blunt nose, however, does not reduce the initialoverpressure rise in the resulting boom signature. Furthermore, a bluntnose creates a significant amount of drag on the aircraft, drasticallydecreasing its efficiency.

U.S. Pat. No. 5,740,984, issued to Morgenstern, for a “Low Sonic BoomShock Control/Alleviation Surfaces” describes a mechanical device on thenose of the airplane which can be moved between a first positioneffecting a blunt nose when sonic boom reduction is desired and a secondposition effecting a streamlined nose when sonic boom reduction is notrequired, thereby removing (in the streamlined configuration) the dragpenalty inherent in a blunt nose design.

U.S. Pat. Nos. 5,358,156, 5,676,333, and 5,251,846, all issued toRethorst and all entitled “Supersonic Aircraft Shock Wave EnergyRecovery System” (collectively “the Rethorst patents”), describe anaircraft with a modified wing design and a forward ring on the fuselagefor eliminating the sonic boom of a supersonic aircraft. FIG. 19 in eachof the Rethorst patents shows a side elevation view of an aircraft whosenose coincides with the bottom of its fuselage. It appears from FIGS.19A and 19B that the bottom of at least a portion of the fuselage isplanar. The Rethorst patents do not provide further disclosure regardingthis fuselage shape, and they do not teach non-uniform propagation ofpressure disturbances about the fuselage. To the contrary, the Rethorstpatents teach that the initial bow shock is axisymmetric about the nose.See U.S. Pat. No. 5,676,333 at col. 14, lines 31-34; U.S. Pat. No.5,738,156 at col. 14, lines 6-10; and U.S. Pat. No. 5,251,846 at col.14, lines 9-12.

Regarding another aspect of the present invention, the same being theinclusion of a leading and/or trailing spike on the supersonic aircraft,the Rethorst patents also describe a supersonic aircraft having a spikeextending from the front of the aircraft and a forward ring on thefuselage for eliminating a sonic boom. The spike is described to directthe bow shock onto the manifold ring that recovers the shock energy andconverts it to useful work. The spike is further depicted as beingextendable, but it does not include a complex surface contour, and it isnot disclosed to include a number of (plurality) telescopicallycollapsible sections. Instead, the Rethorst spike is disclosed as beinga single cylindrical member that tapers to a point at a leading end.

U.S. Pat. No. 4,650,139, issued to Taylor et al., discloses ablunt-nosed spike that can be extended from a space vehicle's fuselage.

U.S. Pat. No. 3,643,901, issued to Patapis, discloses a ducted spike forattachment to a blunt body operating at supersonic speed for the purposeof receiving and diffusing oncoming air to reduce pressure drag on, anderosion of the blunt body.

U.S. Pat. No. 3,425,650, issued to Silva, discloses an apparatus thatcan be extended on a boom from the front of an aircraft to deflect airoutwardly there from.

U.S. Pat. No. 3,655,147, issued to Preuss, covers a device attached tothe lower forebody of an aircraft for the purpose of reflecting pressuredisturbances caused by the aircraft's flight in directions away from theground.

Although some of the foregoing documents are directed to sonic boommitigation, none of them address the sonic boom signature shapingtechniques of the present invention.

DISCLOSURE OF INVENTION

In one embodiment, the invention takes the form of a method forconfiguring and operating an aircraft for minimizing sonic boom effectsat ground level during supersonic flight of the aircraft. The methodincludes configuring the aircraft so that in flight, with landing gearretracted, a lower profile of the aircraft is substantially linear. In arelated embodiment, the profile is slightly concave downward. In eitherembodiment, a nose portion of the aircraft is arranged so that an apexthereof is coincident with the lower profile of the aircraft. Theaircraft is flown at supersonic speed and oriented during supersonicflight so that the substantially linear lower profile of the aircraft isoriented substantially parallel to onset or local airflow. Multipledifferent-magnitude pressure disturbances are generated below theaircraft, and waves thereof are radiated below the aircraft toward theground. These disturbances below the aircraft are of lesser magnitudethan pressure disturbances simultaneously generated and radiated abovethe aircraft. The different-magnitude pressure disturbances generatedbelow the aircraft are controlled so that differentials there among aresufficiently minimized that ground level sonic boom effects areminimized during supersonic flight.

The present invention may be alternatively characterized as a method forconfiguring and operating an aircraft for minimizing sonic boom effectsat ground level during supersonic flight that include configuring theaircraft so that an apex of a nose portion of the aircraft is coincidentwith a lower profile of the aircraft, and when flying the aircraft atsupersonic speed, a majority of a plurality of generateddifferent-magnitude pressure disturbances, and especially the strongestof the generated pressures, are diverted above the aircraft therebyestablishing an asymmetrical distribution of the different-magnitudepressure disturbances about the aircraft. A minority of the plurality ofdifferent-magnitude pressure disturbances that are diverted below theaircraft, and which advantageously constitute the weaker of thedisturbances, are controlled so that ground level sonic boom effects areminimized during supersonic flight.

In a further sense, the present invention(s) relate to aircraft fuselageconfigurations that cause the shock waves created by an aircraft insupersonic flight to propagate non-uniformly about the aircraft suchthat the portions of the shock waves that propagate toward the groundare of lesser intensity than the corresponding portions of the shockwaves produced by an aircraft having a conventional fuselage design. Theamplitude of the sonic boom experienced at the ground is therebyreduced.

A conventional supersonic aircraft includes a generally cylindricalfuselage whose nose comes to a point generally about the fuselage'slongitudinal axis. When such an aircraft flies at supersonic speed, itgenerates shock waves that propagate generally symmetrically in allradial directions about the fuselage.

In the preferred embodiment of the present invention, an aircraftincludes a fuselage whose nose coincides with the bottom of thefuselage. When an aircraft embodying this design flies at supersonicspeed, it creates an asymmetrical pressure distribution. The shock wavesresulting from normal supersonic flight propagate toward the ground withlesser intensity than in other directions. Detailed computational fluiddynamics (CFD) calculations and propagation analyses have shown that asupersonic aircraft embodying the invention produces acharacteristically weaker acoustic signature at the ground than aconventional supersonic aircraft. Thus, the invention provides animportant ingredient for shaping the sonic boom signature to permitsupersonic flight over land.

In another aspect, the present invention provides an additionalimprovement in aircraft design that is directed to mitigating theeffects of sonic booms at ground level. An aircraft according to thepresent invention includes a spike that extends from the aircraft's nosein a direction substantially parallel to the aircraft's length toeffectively lengthen the aircraft. A longer aircraft generally isexpected to produce a sonic boom of lesser amplitude at ground levelthan a shorter aircraft of similar weight because the pressuredisturbance is distributed over a greater length. Therefore, a sonicboom created by an aircraft accordingly configured will be of lesserintensity than a sonic boom created by a conventionally designedsupersonic aircraft having similar characteristics.

The spike can include several sections of varying cross-sectional area.The foremost, or farthest upstream section of the spike preferably has across-sectional area that is characteristically small compared to thatof the aircraft's full fuselage or fuselage forebody. Generally,subsequent (farther aft) downstream sections of the spike progressivelyincrease in cross-sectional area. It is, however contemplated, that aparticular downstream section can have a smaller cross-sectional areathan one or more upstream sections.

Transitions between sections of the spike preferably occur throughcurved or generally conical transition surfaces. However, othertransition region contours are possible, as well. The foremost portionof the spike preferably tapers to a relatively sharp tip at its leadingend, as well as through curved, conical, or other shaped transitionalregions.

In preferred embodiments, the spike can be retracted into the fuselagewhen sonic boom mitigation is not needed or desired. For example, it maybe desirable to retract the spike into the fuselage when the aircraft isflying at subsonic speeds, or is on the ground (to facilitate taxiingand parking).

The spike can be a single member, however it preferably includes two ormore sections that can be collapsed telescopically to facilitateretraction of the spike into the fuselage. Such a telescoping featurealso facilitates adjustment of the spike's overall length and therelative position of the transitions between multiple sections ofvarying cross-sectional area. For example, in the illustrated andexemplary embodiment, the spike includes a substantially cylindricalcenter section (which is the foremost section of the spike when thespike is fully or partially extended) surrounded by one or moreoverlapping, collapsible, annular sections. In other embodiments, theseveral sections can have other regular or irregular cross-sectionalshapes. In such alternate embodiments, the spike can be a single memberor it can be configured as two or more collapsible sections in a mannersimilar to that described above.

When an aircraft embodying such a spike is flown at supersonic speed,the tip of the spike causes an initial shock wave to be formed. Becauseat least the foremost portion of the spike's cross-section ischaracteristically smaller than that of the full fuselage or fuselageforebody, this induced initial shock is of substantially weaker strengththan the initial shock that would be generated by an otherwise unadaptedfuselage or fuselage forebody of an otherwise similar aircraft nothaving a spike. Further weak shocks are caused by the cross-sectionalarea transitions between adjacent telescoping sections (or similardiscontinuities in a one-piece spike's contour), as discussed above.

The position and shape of the foregoing transition regions define thestrength and position of the weak shock waves created thereby. Theposition and shape of these transition regions are selected to reducecoalescence of the weak shocks into a strong sonic boom at the ground.The optimum position and shape of these transition regions are functionsof several variables and can be expected to vary from aircraft toaircraft, based on the particular aircraft's overall configuration. Forexample, the optimum position and shape of the transition regions maydepend on the aircraft's overall length, weight, fineness ratio, wingplacement, engine placement, empennage design and the like. In someembodiments of this aspect of the present invention, the position ofsuch transition regions relative to each other and/or the aircraft'sfuselage can be adjusted on demand by incrementally extending orretracting particular sections of the spike.

A spike according to the present invention can be used in connectionwith conventional fuselage designs. It also can be used in connectionwith other fuselage designs, for example, but without limitation, afuselage configuration in which the nose of the fuselage lies on a linesubstantially defining the bottom of the fuselage; a characteristic thatalso described herein as an aspect or characteristic of a supersonicaircraft configured in conformance with the teachings of the presentinvention(s). As described herein, when an aircraft embodying thisshaped fuselage design flies at supersonic speed, it creates anasymmetrical pressure distribution. The shock waves created by such anaircraft during normal supersonic flight propagate toward the groundwith lesser intensity than in other directions. Detailed computationalfluid dynamics (CFD) calculations and propagation analyses have shownthat such an aircraft can be expected to produce a characteristicallyweaker acoustic signature at the ground than conventional aircraft.Thus, the foregoing fuselage shaping technique provides an importantingredient for shaping the sonic boom signature to permit supersonicflight over land. In alternate embodiments, at least the forward portionof the spike itself can be shaped in a manner similar to the novelfuselage discussed above. A spike embodying such a configuration causesthe portions of the shock waves that propagate toward the ground to beof lesser intensity than the corresponding portions of the shock wavesproduced by an axisymmetric spike.

Similar benefits can be realized from the placement of a spike asdescribed above at the rear of a supersonic aircraft. Accordingly, thepresent invention can be embodied as an aircraft having a spikeprojecting from the aft fuselage or empennage closure thereof inaddition to or instead of the forward-projecting spike described above.

In any event, the several aspects and disclosed embodiments of thepresent invention(s) that are described hereinabove, are not to betreated as limiting, but instead as examples of ways that theinvention(s) can be implemented, as well as claimed for protection asrecited in the attached claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of an aircraft having an axisymmetricfuselage;

FIG. 2A is a top plan view of the aircraft illustrated in FIG. 1;

FIG. 2B is a front elevation view of the aircraft illustrated in FIG. 1;

FIG. 2C is a side elevation view of the aircraft illustrated in FIG. 1;

FIG. 3A is a perspective view of an aircraft represented as anequivalent body of revolution, including the effects of lift and volume;

FIG. 3B is a front elevation view of an aircraft represented as anequivalent body of revolution, showing the transition from asubstantially cylindrical cross-section to a point;

FIG. 4 illustrates the near-field pressure contour produced by anaircraft represented as an equivalent body of revolution flying atsupersonic speed;

FIG. 5 illustrates the propagation of the pressure disturbance producedby an aircraft represented as an equivalent body of revolution flying atsupersonic speed;

FIG. 6A is a plot of the near-field pressure disturbance caused by anaircraft represented as an equivalent body of revolution traveling atsupersonic speed;

FIG. 6B is a schematic plot of the pressure disturbance at ground levelcaused by an aircraft represented as an equivalent body of revolutiontraveling at supersonic speed;

FIG. 7 is a perspective view of an aircraft having an asymmetricfuselage, at least with respect to horizontal, and configured accordingto the present invention;

FIG. 8A is a top plan view of the aircraft illustrated in FIG. 7;

FIG. 8B is a front elevation view of the aircraft illustrated in FIG. 7;

FIG. 8C is a side elevation view of the aircraft illustrated in FIG. 7;

FIG. 9A is a perspective view of an aircraft represented as anequivalent asymmetric body according to the present invention, includingthe effects of lift and volume;

FIG. 9B is a front elevation view of an aircraft represented as anequivalent asymmetric body according to the present invention, showingthe transition from a substantially cylindrical cross-section to a pointaligned with the bottom of the body;

FIG. 10 illustrates the near-field pressure contour produced by anaircraft represented as an equivalent asymmetric body according to thepresent invention flying at supersonic speed;

FIG. 11 illustrates the propagation of a pressure disturbance producedby an aircraft represented as an equivalent asymmetric body flying atsupersonic speed;

FIG. 12A is a schematic plot of the near-field pressure disturbancecaused by an aircraft represented as an equivalent asymmetric bodyaccording to the present invention traveling at supersonic speed (dashedline), superimposed on a plot of the near-field pressure disturbancecaused by a conventionally designed aircraft represented as anequivalent body of revolution traveling at supersonic speed (solidline);

FIG. 12B is a plot of the pressure disturbance at ground level caused byan aircraft represented as an equivalent asymmetric body according tothe present invention traveling at supersonic speed (dashed line),superimposed on a plot of the pressure disturbance at ground levelcaused by an aircraft represented as an equivalent body of revolutiontraveling at supersonic speed (solid line).

FIG. 13 is a perspective view of an alternative embodiment of thepresent invention in which the lower profile is substantially linear,but not absolutely linear;

FIG. 14 is an elevational view of the aircraft of FIG. 13 shown relativeto the ground and with the propagation of pressure disturbances depicteddown to the ground where a sonic boom effect is minimized;

FIG. 15 is a perspective view of an alternative embodiment of thepresent invention in which the lower profile is downwardly concave;

FIG. 16 graphically demonstrates near-body (near-field) pressuredisturbances generated by a conventionally configured sonic aircraftgenerally in association with the length of the aircraft;

FIG. 17 graphically demonstrates distant or ground-effect pressuredisturbances generated by a conventionally configured sonic aircraftroughly in association with the length of the aircraft and causing sonicboom;

FIG. 18 graphically demonstrates near-body (near-field) pressuredisturbances generated by a sonic aircraft configured according to thepresent invention and generally associated with the length of theaircraft;

FIG. 19 graphically demonstrates distant or ground-effect pressuredisturbances generated by a sonic aircraft configured according to thepresent invention roughly in association with the length of the aircraftin which sonic boom effect has been minimized;

FIG. 20 is a schematic representation of the development of sonic boomby a conventionally designed sonic speed aircraft;

FIG. 21 is a pictorial of an exemplary aircraft in which abelow-fuselage engine creates a disadvantageous pressure disturbance;

FIG. 22 is a pictorial of an exemplary aircraft flying at an inclinedangle of attack;

FIG. 23 is a schematic comparative view showing a supersonic aircraftconfigured according to the teachings of the present invention flying,together with its associated shaped sonic boom signature along side aconventionally designed supersonic aircraft with its generated N-shapedsonic boom signature; and

FIG. 24 is a perspective view of a supersonic aircraft having anelongated spike extending from its nose according to the presentinvention;

FIG. 25 depicts a series of side elevation views of an aircraft noseoutfitted with a telescopically collapsible spike configured accordingto the present invention, illustrating the spike in various degrees oftelescopic extension/retraction; and

FIG. 26 is a plot of the initial pressure rise at ground levelassociated with the bow shock created by a conventional aircraft flyingat supersonic speed superimposed on a plot of the initial pressure riseassociated with the bow shock created by an aircraft outfitted with aspike according to the present invention flying at supersonic speed.

FIG. 27 is an illustration of a non-limiting embodiment of acontinuously tapered spike.

MODE FOR INVENTION

The propagation characteristics of shock waves created by supersonicaircraft can be analyzed using, for example, CFD analysis methods. Theseanalyses can be complicated because an aircraft includes many components(for example, a fuselage, wings, engines, tailfin, etc.) that contributeto such disturbances. However, such analyses commonly are simplified bymodeling the aircraft as a semi-infinite body of revolution. Analysesindicate that shock waves propagate substantially uniformly aboutsupersonic aircraft modeled in this manner.

FIGS. 3A and 3B provide perspective and front elevation views,respectively, of an aircraft represented as a semi-infinite equivalentbody of revolution 22, with the front of the aircraft corresponding topoint 36 on the equivalent body of revolution. Equivalent body ofrevolution 22 models the atmospheric disturbance caused by the flight ofthe aircraft it represents. More particularly, equivalent body ofrevolution 22 models the atmospheric disturbance caused by thedisplacement of atmospheric medium by the volume of the aircraft and bythe lift generated by the aircraft. Portion 37 of equivalent body ofrevolution 22 represents the disturbance caused by such volume and lift,while the remainder of equivalent body of revolution 22 represents thedisturbance caused by lift only. As such, portion 37 of equivalent bodyof revolution 22 corresponds to the length of the aircraft representedthereby, while the remainder of equivalent body of revolution 22corresponds to the wake thereof. As is most clearly illustrated in FIG.3B, each cross-section of equivalent body of revolution 22 issubstantially circular, and the center of each such circularcross-section lies on a common centerline 24.

FIG. 4 illustrates a computer model of the near-field pressuredisturbance that would be created by an aircraft represented asequivalent body of revolution 22 flying at supersonic speed. Thispressure disturbance is characterized by bow shock 26 which propagatessubstantially uniformly, i.e., axisymmetrically, about equivalent bodyof revolution 22 and, thus, the aircraft it represents. Bow shock 26propagates in the shape of a Mach cone, as described above. As shown inFIG. 5, bow shock 26 remains axisymmetric about equivalent body ofrevolution 22 as bow shock 26 propagates far away from the aircraft; thetail shock 27 behaves similarly as shown.

FIG. 6A is a graph of the near-field pressure disturbance 40 (thepressure disturbance near the aircraft) caused by an aircraftrepresented as equivalent body of revolution 22 traveling at supersonicspeed as a function of location relative to the aircraft. The x-axisunits are X°−° Y/tan(μ), where X represents the axial location of apoint on the aircraft measured from the front of the aircraft, Yrepresents the perpendicular distance from the aircraft to the pointwhere the disturbances are being modeled (here, Y is about equal to 2.5times the length of the of the aircraft) and μ is the Mach angle, asexplained above. The y-axis units are Δ P/P, where P represents ambientpressure and Δ P represents the change in local pressure from ambientpressure.

The near-field pressure disturbance is characterized by a positivepressure spike 42 occurring at about the nose of an aircraft representedas equivalent body of revolution 22, followed by a sharp pressurereduction 44 between the nose and tail of such an aircraft to belowambient pressure, followed by a gradual return to ambient pressure 46 atabout the tail of such an aircraft.

At greater distances Y from an aircraft represented by equivalent bodyof revolution 22, the individual pressure waves contributing to thenear-field distribution illustrated in FIG. 6A coalesce to form aclassic sonic boom acoustic signature, or N-wave, 50 as shownschematically in FIG. 6B, wherein the value of Y (i.e., theperpendicular distance from the aircraft to the point where thedisturbance is being measured) is taken to be about 500 times the lengthof the aircraft. The acoustic signature 50 of an aircraft represented asequivalent body of revolution 22, shown schematically in FIG. 6B, ischaracterized by a positive pressure spike 52 corresponding to the bowshock passing a reference point (e.g., a point on the ground), followedby a linear pressure decrease to sub-ambient pressure 54, followed by asecond positive pressure spike 56 corresponding to the tail shockpassing the reference point, returning the pressure to ambient pressure.

FIG. 1 provides a perspective view of a conventional aircraft 20, whichcan be readily represented by equivalent body of revolution 22, as shownin FIGS. 3A and 3B. Aircraft 20 includes wings 28 and engines 34attached to a substantially axisymmetric fuselage 21. Aircraft 20further includes horizontal stabilizer 32 and tailfin 30, both of whichin turn are attached to fuselage 21. FIGS. 2A-2C provide top plan, frontelevation, and side elevation views, respectively, of conventionalaircraft 20.

FIGS. 9A and 9B illustrate perspective and front elevation views of anaircraft configured according to the present invention that isrepresented as equivalent body 122. Equivalent body 122 models theatmospheric disturbance caused by the flight of aircraft according tothe present invention. More particularly, equivalent body 122 models theatmospheric disturbance caused by the displacement of atmospheric mediumby the volume of an aircraft according to the present invention and bythe lift generated by such an aircraft. Portion 137 of equivalent body122 represents the disturbance caused by such volume and lift, while theremainder of equivalent body 122 represents the disturbance caused bylift only. As such, portion 137 of equivalent body 122 corresponds tothe length of the aircraft represented thereby, while the remainder ofequivalent body 122 corresponds to the wake thereof.

It can be seen from FIGS. 9A and 9B that equivalent body 122 is not abody of revolution, but is instead asymmetric. These figures,particularly FIG. 9B, further show that each cross-section of equivalentbody 122 may be substantially circular in the preferred embodiment.However, whereas the centers of each cross-section of equivalent body ofrevolution 22 illustrated in, for example, FIGS. 3A and 3B, lie on acommon centerline 24, the same is not true of the cross-sections ofequivalent body 122. Instead, the bottom of substantially each and everycircular cross-section of equivalent body 122 lies substantially on acommon line 124. As will be discussed further below, the bottom of atleast a substantial portion of the cross-sections comprising at leastthe forward portion of an aircraft fuselage according to the presentinvention; i.e., an aircraft represented by equivalent body 122, lie ona common line.

FIG. 10 illustrates a computer model of the near-field pressuredisturbance that would be created by an aircraft represented byequivalent body 122 flying at supersonic speed. Like the near-fieldpressure disturbance caused by equivalent body of revolution 22,illustrated in FIG. 4, these pressure disturbances are characterized bybow shock 126 that propagates about equivalent body 122 in the shape ofa Mach cone and tail shock 127 as shown. However, the pressuredisturbance caused by equivalent body 122 is markedly different from thepressure disturbance caused by equivalent body of revolution 22 in thatthe pressure contour associated with the disturbance caused byequivalent body 122 is much stronger above and to the sides thereof thanbeneath it. That is, the pressure contour associated with thisdisturbance is asymmetric. Further, the pressure contour beneathequivalent body 122 is much less dense than the pressure contour beneathequivalent body of revolution 22, representing a conventional aircraftof similar size, under similar flight conditions. As shown in FIG. 11,the pressure contour resulting from bow shock 126 remains asymmetricabout equivalent body 122 as bow shock 126 propagates away fromequivalent body 122.

FIG. 12A provides a graph of the near-field (here, Y is about equal to2.5 times the aircraft length) pressure disturbance 140 caused by anaircraft represented by equivalent body 122 traveling at supersonicspeed, superimposed on the graph of the near-field pressure disturbance40 caused by an aircraft represented by equivalent body of revolution 22traveling at supersonic speed, as illustrated in FIG. 6A. The peakpressure rise 142 resulting from supersonic flight of an aircraftrepresented by equivalent body 122 is of substantially lesser magnitudethan the peak pressure rise 42 caused by an aircraft represented byequivalent body of revolution 22 under similar flight conditions.Similarly, the pressure drop 144 to below ambient associated with anaircraft represented by equivalent body 122 is of substantially lessermagnitude than pressure drop 44 to below ambient caused by an aircraftrepresented by equivalent body of revolution 22 under similar flightconditions. Likewise, the pressure return 146 to ambient associated withan aircraft represented by equivalent body 122 is of lesser magnitudethan pressure return 46 to ambient caused by an aircraft of similar sizerepresented by equivalent body of revolution 22, under similar flightconditions.

FIG. 12B provides a graph of the far-field (here, Y is about equal to500 times the aircraft length) pressure disturbance 150 caused by anaircraft according to the present invention represented by equivalentbody 122 traveling at supersonic speed, superimposed on the graph of thefar-field pressure disturbance 50 caused by an aircraft represented byequivalent body of revolution 22 traveling at supersonic speed, asillustrated in FIG. 6A. The peak pressure rise 152 resulting fromsupersonic flight of an aircraft represented by equivalent body 122 isof substantially lesser magnitude than the peak pressure rise 52 causedby an aircraft of similar size represented by equivalent body ofrevolution 22, under similar flight conditions. Similarly, the pressuredrop to below ambient 154 associated with an aircraft represented byequivalent body 122 is of substantially lesser magnitude than pressuredrop 54 to below ambient caused by an aircraft represented by equivalentbody of revolution 22 under similar flight conditions. Likewise, thepressure return to ambient 156 associated with an aircraft representedby equivalent body 122 is of substantially lesser magnitude thanpressure return 56 to ambient caused by an aircraft represented byequivalent body of revolution 22 under similar flight conditions.

CFD analysis thus shows that the pressure disturbance above an aircraftconfigured according to the present invention represented by equivalentbody 122 is significantly greater than the pressure disturbance belowsuch an aircraft. Relatively strong disturbances, shown as tightlypacked contour lines in FIGS. 10 and 11, propagate upward, away from theground. Substantially weaker disturbances, shown as loosely packedcontour lines in FIGS. 10 and 11, propagate towards the ground. Further,the ground-directed disturbances produced by an aircraft represented byequivalent body 122 are substantially weaker than the ground-directeddisturbances produced by an aircraft represented by equivalent body ofrevolution 22. Thus, the ground-directed disturbances produced by anaircraft represented by equivalent body 122 according to the presentinvention are expected to result in significantly weaker sonic boomscompared to those produced by an aircraft represented by equivalent bodyof revolution 22.

FIG. 7 illustrates a perspective view of a supersonic aircraft 120having a novel fuselage design according to a preferred embodiment ofthe present invention. FIGS. 8A-8C illustrate top plan, front elevation,and side elevation views of aircraft 120, respectively. The foregoingfigures illustrate a preferred embodiment of the invention wherein thebottom of substantially every cross-section of fuselage 121 liessubstantially on a line located at the intersection of the bottom offuselage 121 with a plane tangent to the bottom of fuselage 121, asdescribed above. In certain alternate embodiments, many of the benefitsof the foregoing fuselage design can be realized even if the bottom ofsome cross-sections of fuselage 121 do not lie on such a line. Forexample, in one alternate embodiment (not shown), fuselage 121 isasymmetric at its nose, but axisymmetric at its tail. In thisembodiment, the bow shock experienced at ground level is of lessermagnitude than the bow shock resulting from supersonic flight of anaircraft having an axisymmetric nose. Other alternate embodiments mayinclude discontinuities in the configuration of the fuselage bottom suchthat some cross-sections of the fuselage do not include a point thatlies on a line formed by the intersection of the bottom of the fuselageand a plane tangent thereto. In fact, physical limitations associatedwith aircraft construction may preclude a configuration wherein thebottom of each and every fuselage cross-section lies on such a line,although such a configuration is within the scope of the presentinvention.

Further, although fuselage 121 is shown in FIGS. 7 and 8A-8C as havingsubstantially circular cross-sections, fuselage 121 could have differentcross-sectional shapes (or combinations of cross-sectional shapes) inother embodiments. Examples of such other cross-sectional shapesinclude, without limitation, non-circular curved shapes, partiallycircular shapes, partially non-circular curved shapes, and angled shapes(e.g., a “V” shape). Further, a fuselage according to the presentinvention can include more than one of the foregoing (or other)cross-sectional shapes along its length.

In a preferred embodiment, the invention takes the form of a method forconfiguring and operating an aircraft for minimizing sonic boom effects172 at ground level during supersonic flight of the aircraft. Oneexample of such a preferred embodiment is shown in the perspective viewsof FIGS. 7 and 13 where an airplane 120 is shown flying at supersonicspeed without creating a conventional sonic boom at ground level. Thisminimization of sonic boom signature is attributable at least in part tothe fact that the craft is configured, so that in flight, and withlanding gear retracted, the presented lower profile 160 of the aircraft120 is substantially linear in configuration. To this end, a noseportion 162 of the fuselage 121 of the aircraft 120 is arranged so thatan apex 163 thereof is coincident with the lower profile 160 of theaircraft. Exemplarily, it is the lower exterior surface 164 of thefuselage 121 of the aircraft 120 that establishes this substantiallylinear lower profile 160 of the aircraft 120.

This embodiment of the invention includes not only this structuralconfiguration of the aircraft 120, but also flying the aircraft 120 atsupersonic speed and orienting the aircraft 120 during such supersonicflight so that the lower profile 160 of the craft 120 is orientedsubstantially parallel to onset airflow 166. Onset flow 166 isillustrated in FIGS. 7 and 13, among others, by the arrow located aheadof the craft 120, and which is pointing toward the nose 163 of the craft120. This onset flow 166 may be thought of as the relationship betweenthe craft 120 and the air that is flowing there over. In actually,however, it is the relative orientation of the airplane 120 as it pushesthrough the air. As explained herein, as such an aircraft 120 pushesthrough the air at supersonic speeds, pressure disturbances or waves areproduced thereabout. As an aspect of the present invention, multiple, oras otherwise referred to, a plurality of different-magnitude pressuredisturbances 168 are generated below the aircraft 120 and which thenradiate there below. Conceptually, these pressure disturbances areillustrated in FIGS. 10 and 11. These aspects are graphically shown inFIGS. 12A, 12B, and 16-19 by dashed lines, and comparison is made incertain of these drawings to conventional characteristics oftraditionally configured supersonic aircraft which are represented bysolid line traces. These generated disturbances below the craft 120 areof lesser magnitude than a plurality of pressure disturbances 170simultaneously generated above the aircraft 120 and radiating thereabove. An important feature of this embodiment of the invention is thatthe structural design of the craft 120 enables this plurality ofdifferent-magnitude pressure disturbances 168 generated below theaircraft to be controlled so that differentials there among (across theseveral pressure disturbances) are sufficiently minimized that groundlevel sonic boom effects are minimized during supersonic flight.

Throughout the description of the invention, certain aspects arecharacterized with the qualifier “substantially.” For interpretationpurposes, this terminology should be taken to denote the fact thatmoderate variations may be made from the so described configuration,orientation or relationship, but within limits that continue theprescribed effects associated with the so described aspect.

An aspect of the above-described embodiment of the invention is thatduring supersonic flight, the aircraft 120 is preferably oriented sothat its lower substantially linear profile is leveled to besubstantially parallel with the direction of travel and onset airflow166. This orientation is illustrated in at least FIG. 14, and can becompared to more traditional flying configurations such as that shown inFIG. 22 where a wing reference plane is shown flying with an inclinedangle of attack. It should be appreciated that such an inclined angle ofattack tends to accentuate downwardly directed pressure disturbances, asopposed to minimizing them as is the case in the more horizontal flyingorientation of the present invention. Regarding illustrations in theassociated drawings in which supersonic aircraft are shown relative tothe ground, it should be appreciated that these FIGS. are not to scale,especially with respect to the perception of the elevations at which theaircraft are flying.

Another aspect of the invention is that the aircraft 120 is configuredso that during supersonic flight coalescence of the different-magnitudepressure disturbances 168 is at least inhibited below the aircraft 120.This can be compared to traditional effects such as illustrated in FIG.20 where bow shocks are shown to coalesce into a strong overpressure andthe tail shocks are shown to coalesce into a drastic return towardambient pressure. These coalescing effects contribute to the formationof traditional “N-wave” sonic boom signatures experienced at groundlevel below conventionally designed and flown supersonic aircraft.

In a particularly preferred aspect of the invention, and as illustratedin at least FIG. 14, the lower profile 160 of the aircraft 120 isconfigured so that none of the stronger pressure disturbances generatedbelow the aircraft and behind the disturbance caused by the forward nose163 of the craft 120 propagate at angles sufficient to result in theircoalescence prior to reaching ground level. This is schematicallyrepresented by the pressure lines stemming from features of the aircraft120 behind the nose tip 163 in FIG. 14.

In the present context, the terminology of stronger pressuredisturbances is used to identify those pressure disturbances ofsufficient magnitude to have a potential for coalescing withdisturbances in front or behind thereof, and thereby combining into asingle, stronger disturbance instead of the previously distinct, weakerones.

A preferred embodiment of the present invention also shapes the noseportion 162 of the aircraft 120 so that vertical cross-sections orientedperpendicular to a long axis of the aircraft are substantiallyround-shaped. Schematically, this is shown in FIG. 9B where a lowerextremity of each circle 122 is coincident with the lower profile 160(124) of the fuselage 121.

It is also contemplated that these vertical cross-sections can besubstantially elliptical-shaped, with long axis being eithersubstantially vertical or substantially horizontal.

In another aspect of the presently disclosed invention(s), the jetpropulsion units 134 mounted upon the aircraft 120 are configured sothat resulting pressure disturbances 168 created thereby and below theaircraft 120 are of lesser magnitude than any pressure disturbancecaused by the apex 163 of the nose portion 162 below the aircraft 120.Still further, all inlets 135 of side-mounted jet propulsion units arepositioned at above-wing locations thereby assuring that downwardlydirected pressure disturbances 168 generated by the inlets 135 aresubstantially blocked from direct propagation below the substantiallylinear lower profile 160 of the aircraft 120.

A primary method of the presently disclosed invention for controllingthe plurality of different-magnitude pressure disturbances 168 generatedbelow the aircraft 120 is by selective arrangement of discontinuities165 in a lower exterior surface of the aircraft and thereby assuringthat ground level sonic boom effects are minimized during supersonicflight. Discontinuities should be understood to be established by slopechanges across features of the aircraft 120 that establish the profilethereof.

In FIG. 15, an alternative embodiment of the present invention isillustrated in which the lower profile 160 of the aircraft 120 isuniquely mildly downwardly concave. As in FIG. 14, none of the strongerpressure disturbances generated below the aircraft and behind thedisturbance caused by the forward nose 163 of the craft 120 propagate atangles sufficient to result in their coalescence prior to reachingground level.

Another way of characterizing the present invention is that afterconfiguring an apex 163 of a nose portion 162 of the aircraft 120 to becoincident with a lower profile 160 of the aircraft 120, the aircraft120 is flown at supersonic speed and a majority of the generateddifferent-magnitude pressure disturbances 168 are diverted above theaircraft 120 thereby establishing an asymmetrical distribution of thedifferent-magnitude pressure disturbances 168, 170 thereabout. Inconjunction therewith, a minority of the plurality ofdifferent-magnitude pressure disturbances 168 that are diverted belowthe aircraft 120 are controlled so that ground level sonic boom effectsare minimized during supersonic flight.

In another aspect, the invention takes the form of a method forminimizing sonic boom effects caused at ground level by a supersonicaircraft. The method includes manipulating at least one sonic boomcontributing characteristic of a supersonic aircraft to assure that aplurality of groundwardly radiating pressure disturbances do notcoalesce, one with another, to form an objectionable sonic boom duringsupersonic flight by the aircraft.

A related characterization of the invention entails manipulating atleast one sonic boom contributing design characteristic of a supersonicaircraft to prevent coalescence of groundwardly radiating pressuredisturbances, generated during supersonic flight, and therebyestablishing a shaped sonic boom signature 180 of the aircraft, atground level, that is humanly perceptible, but non-objectionable to aperceiving person located on the ground. An example of such a designcharacteristic is found in the aspect described herein regarding theconfiguration of a lower profile of the supersonic aircraft so that anapex of a nose portion of the aircraft is coincident with a lowerprofile of the aircraft. This embodiment of the method further includesflying the aircraft at supersonic speed and diverting a majority of aplurality of generated different-magnitude pressure disturbances abovethe aircraft thereby establishing an asymmetrical distribution of thedifferent-magnitude pressure disturbances about the aircraft such thatthe objectionable ground level sonic boom effects are minimized.

In this regard, FIG. 23 shows a comparison between a conventionallydesigned supersonic aircraft 20 at the left, including its N-shapedsonic boom signature 50 which is unacceptable to persons located atground level. On the right, an aircraft 120 configured according to theexemplary embodiment described immediately above, and which produces anon-offending shaped sonic boom signature 180 at ground level, isillustrated.

It should be appreciated that presently regulations generally preventcivil supersonic flight over land. Studies conducted with humanparticipants, however, show that sonic boom effects, at ground level, inand of themselves are not always found to be objectionable by a humanreceiver. Sonic boom effects are only bothersome to humans located onthe ground when they are sufficiently loud and abrupt (strong ΔP andshort rise time to peak overpressures) to be objectionable. A parallelmay be drawn to noise level regulations instituted with respect toairports. That is to say, take-off noise levels are limited, notprecluded by such regulations. Therefore, it is in this vein that theterminology used in characterizing the present invention is found;namely, that a shaped sonic boom signature 180 is established, viamanipulation of characteristic(s) of a supersonic aircraft thatinfluence sonic boom effects imposed at ground level, but with thequalifier that they be humanly perceptible and non-objectionable to aperceiving person located on the ground. Studies that quantify suchsonic boom effects that are, and are not objectionable to people areknown to those persons skilled in these arts, and therefore may bereadily applied, from a definitional standpoint, to such recitationsfound herein.

In a second aspect, a supersonic aircraft can be configured to include aspike extending from the front thereof. For example, FIG. 24 illustratesa supersonic aircraft 220 having a spike 223 extending forward fromfuselage 221, generally in the direction of normal flight. Fuselage 221can be otherwise conventional, similar to fuselage 21 described above,or it can be specially shaped, similar to fuselage 121 also describedhereinabove. Alternatively, fuselage 221 can have other configurations.

Spike 223 preferably can be at least partially retracted into thefuselage of the aircraft on demand. For example, it may be desirable toretract spike 223 into fuselage 221 when the aircraft 220 is flown atsubsonic speeds, flown at supersonic speed over areas where sonic boomsare deemed acceptable (such as over an ocean), and/or on the ground (tofacilitate taxiing and parking).

In a preferred embodiment, spike 223 has a forward section 223A and arearward section 223B. With reference to FIG. 25, forward section 223Ahas a generally smaller nominal cross-sectional area than does rearwardsection 223B, which, in turn, has a generally smaller nominalcross-sectional area than does fuselage 221. Forward section 223A taperstoward (i.e., to, or substantially to) a point 223C through transitionregion 223D. In alternate embodiments, forward section 223A can tapertoward other shapes. For example, but without limitation, forwardsection 223A can taper toward an edge, such as a knife-edge, which canbe oriented vertically, horizontally, or in any other desirable manner.

The transition from forward section 223A to rearward section 223B isthrough transition region 223E. Transition region 223D is shown assubstantially conical and transition region 223E is shown assubstantially frusto-conical. These transition regions, however, canhave curved or other contours as well. In other configurations of thisaspect of the invention, spike 223 can have one or more additionalsections between rearward section 223B and fuselage 221. An additionaltransition region, as discussed above, would be associated with eachsuch additional section. Generally, the nominal cross-sectional area ofany such additional section would be greater than the nominalcross-sectional area of a section forward thereof, and smaller than thatof a section rearward thereof. However, it is possible that such anintermediate section could have a nominal cross-sectional area smallerthan that of a section forward thereof and/or larger than that of asection rearward thereof. Generally, the nominal cross-sectional area ofany section of spike 223 is substantially smaller than the nominal crosssectional area of fuselage 221. Although the nominal cross-sectionalarea of each section of spike 223 is shown to be substantially uniformover the length thereof, the cross-sectional area of each section canvary over the length thereof.

FIGS. 24 and 25 illustrate spike 223 as having substantially cylindricalcross-sections. In other embodiments, it is contemplated that spike 223can have other regularly or irregularly shaped cross-sections.

Spike 223 can be embodied as a single member. However, it is preferredthat sections 223A and 223B (as well as any additional sections, asdiscussed above) be separate elements which are collapsible in atelescoping manner. FIG. 25 shows a preferred embodiment of atelescopically collapsible spike 223 in an extended position A, aretracted position D, and two intermediate positions B and C.

In alternative embodiments, the spike 223 could be of a single, taperedsection. Alternatively, spike 223 can have several sections, one or moreof which are tapered continuously over the length thereof. The severalsections can be collapsible, or embodied as a single member. An exampleof a continuously tapered spike 323 is illustrated in FIG. 27 andincludes a first section 323A and s second section 323B. A transitionregion between the trailing end of spike 323 and a leading end offuselage 221 is identified by the reference letter E.

When an aircraft 220 that includes a spike 223 as illustrated in FIGS.24 and 25 is flown at supersonic speed, the tip of the spike causes aninitial shock wave to be formed. Because the spike's cross-section(taken in a generally perpendicular orientation to a long axis of theaircraft 220), is substantially smaller than that of the aircraft's fullfuselage or fuselage forebody, this initial shock is substantiallyweaker than the initial shock that would be created by the full fuselageor fuselage forebody of an otherwise similar aircraft not having aspike. The initial shock on the spike is also well in front of the shockcaused by the fuselage forebody and therefore the spike is bothweakening the initial shocks and also lengthening the sonic boomsignature that is propagated to the ground. A further weak shock iscaused by each further transition region (such as transition region223E) between adjacent sections (such as sections 223A and 223B) ofspike 223. As the number of sections of spike 223 increases, the numberof transition regions increases, and the number of weak shocks createdthereby increases.

The position and shape of the transition regions define the strength andposition of the weak shocks created thereby. The position and shape ofthese transition regions are selected to reduce coalescence of the weakshocks into a strong shock and thus reduce the intensity of a sonic boomat ground level resulting from these shocks. As discussed above, theoptimum position and shape of these transition regions are functions ofseveral variables and can be expected to vary from aircraft to aircraft,based on the particular aircraft's overall configuration. For example,the optimum position and shape of the transition regions may depend onthe aircraft's overall length, weight, fineness ratio, wing placement,engine placement, empennage design, altitude, Mach number (speed) andrelated characteristics. In some embodiments of this aspect of thepresent invention, the position of such transition regions relative toeach other and/or the aircraft's fuselage can be adjusted on demand byincrementally extending or retracting particular sections of the spike.For example, referring to FIG. 25, it may be desirable under certaincircumstances to operate the aircraft with spike 223 in position B,position C, or another intermediate position.

FIG. 26 illustrates graphically the effect of spike 223 on the shockcreated by an aircraft equipped therewith during supersonic flight. FIG.26 provides a plot 230 of the pressure rise associated with the bowshock created by an aircraft flying at supersonic speed that has beenadapted to project a shaped signature to the ground as described herein,superimposed on a plot 240 of the pressure rise associated with the bowshock created by a similar aircraft having a spike 223 in an extendedposition and flying at supersonic speed. FIG. 26 shows that an aircraft220 having such a spike 223 and flown at supersonic speed produces asubstantially lower initial pressure rise 242 than the initial pressurerise 232 created by a conventional aircraft of similar size undersimilar flight conditions. Further, the peak pressure rise resultingfrom supersonic flight of aircraft 220 having spike 223 is reachedthrough a series of relatively small step increases in pressure 242,244, 246, 248, whereas the peak pressure rise resulting from supersonicflight of conventional aircraft 220 is reached through a series offewer, but larger, step increases in pressure 232,234,236 (notnecessarily shown to exact scale in FIG. 26). Generally, the sonic boomat ground level will be reduced where the peak pressure rise is realizedthrough a longer series of smaller pressure increases, instead ofthrough a shorter series of larger pressure increases.

It should also be appreciated that spike 223 can be used in connectionwith otherwise conventional supersonic aircraft 20 to effect a reductionin the sonic boom experienced at ground level. Spike 223 also can beused in connection with supersonic aircraft having a specially shapedfuselage 121 as described hereinabove. In certain contemplatedconfigurations, spike 223, itself, can be specially shaped in a mannersimilar to that of shaped-fuselage 121.

An aircraft according to the present invention can have a second spikesimilar to spike 223 extending from the aft fuselage or empennageclosure thereof in addition to spike 223 extending from the forwardfuselage thereof. In alternate embodiments, such an aircraft can havesuch a rearwardly projecting spike instead of a forward projecting spike223.

While the foregoing embodiments of the invention illustrate a supersonicpassenger jet, it should be understood that the configuration can beused in connection with other types of aircraft and aerospace vehicles.

Whereas the present invention is described herein with respect tospecific embodiments thereof, it will be understood that various changesand modifications may be made by one skilled in the art withoutdeparting from the scope of the invention, and it is intended that theinvention encompass such changes and modifications as fall within thescope of the appended claims.

INDUSTRIAL APPLICABILITY

The present invention finds industrial applicability at least within thesupersonic categories of aircraft and aerospace industries.

We claim:
 1. An aerospace vehicle configured to reduce the effects of asonic boom created by said vehicle when said vehicle is flown atsupersonic speed, said aerospace vehicle comprising: a fuselage having aleading end, a trailing end, a top, and a bottom; a spike operablyassociated with the leading end of said fuselage, the spike having aspike leading end and a spike trailing end, the spike leading endtapering toward the spike trailing end, a cross-sectional magnitude ofthe spike not decreasing from the spike leading end to the spiketrailing end, the spike leading end comprising a forward-most portion ofthe aerospace vehicle, the spike having a plurality of sections, and thespike being substantially free of outwardly extending projections; and atransition region between the spike trailing end and the leading end ofthe fuselage, the transition region comprising a substantially smoothand gradual change at the spike trailing end from a first contour to asecond contour, the first contour substantially conforming to a trailingportion of the spike and the second contour substantially conforming tothe leading end of the fuselage, and wherein a longitudinally centralregion of each section of the spike has a substantially uniformdiameter.
 2. The aerospace vehicle of claim 1, wherein the spikecomprises a first section aft of the spike leading end and a secondsection aft of the first section.
 3. The aerospace vehicle of claim 2,wherein the first section includes a first section leading end and afirst section trailing end and the second section has a second sectionleading end and a second section trailing end.
 4. The aerospace vehicleof claim 3, wherein the second section trailing end is located in thetransition region of the spike.
 5. The aerospace vehicle of claim 4,wherein one of the first and second sections is selectively collapsibleinto the other of the first and second sections.
 6. The aerospacevehicle of claim 1, wherein the spike is selectively extendable from theleading end of the fuselage.
 7. The aerospace vehicle of claim 1,wherein the spike is selectively retractable into the leading end of thefuselage.
 8. The aerospace vehicle of claim 1, wherein the spike leadingend tapers in a forward direction toward a point.